The present invention relates to actively-cooled tubular shell structures for high-temperature applications having a plurality of cooling channels formed between the inner wall and outer wall of the structure wherein the structure is fabricated from fiber-reinforced ceramic matrix composite materials and methods of producing same.
Performance of advanced chemical rocket technology is limited, for the most part, by the availability of high-temperature structural engineering materials. Greater performance and efficiency in liquid rocket propulsion systems can be gained by operating at higher combustion temperatures and higher working pressures resulting from highly energetic propellant mixture ratios (oxidizer/fuel). However, propellant mixtures are typically xe2x80x9ctamedxe2x80x9d to off-optimum conditions in order not to exceed the temperature limitations of the thrust chamber and nozzle materials.
A considerable amount of heat is transferred in all designs of rocket engines. The principle objective of high-temperature rocket design is to safely limit the heat transfer to the materials in critical hot sections such as the injector, combustion chamber, throat, and nozzle. A failure would impair the satisfactory operation of the rocket propulsion system or the flight vehicle being propelled. The walls have to be maintained sufficiently cool so that wall temperatures do not exceed their safe allowable operating limit. Erosion, usually the result of combined oxidation and chemical interaction with the hot combustion gases, should not damage the walls, and the walls should be capable of withstanding the extreme thermal shock caused by the sudden onset of a high heat flux from combustion ignition. The materials comprising the thrust chamber devices must also be capable of resisting the thermal stresses induced by the heat transfer and thermal gradients.
There are two general cooling methods commonly used today in the design of liquid propellant rocket engine devices. Those devices that reach thermal equilibrium during operation typically operate for long durations and are usually either actively-cooled (e.g., bipropellants) or radiation cooled (e.g., monopropellants).
Actively-cooled liquid propellant thrust chambers have provisions for cooling some or all of the components in contact with the hot combustion gases, such as the chamber walls, nozzle walls and injector faces. A cooling jacket or cooling coil often consists of separate inner and outer walls or a bundled assembly of continuous, contoured tubes. The inner wall confines the combustion gases, and the space between the inner and outer walls serves as the coolant passage. The axial or helical passages in liquid propellant rocket thrust chambers are often of complex cross-section. The nozzle throat region is usually the location that sustains the greatest heat transfer intensity and is therefore the most difficult to cool. For this reason the cooling jacket is often designed so that the coolant velocity is highest at the critical regions by restricting the coolant passage cross-section and so that the coolant enters the jacket at or near the nozzle.
Regenerative cooling is a form of active cooling and is used for engines where one of the propellant constituents is circulated through cooling passages around the thrust chamber prior to injection and burning of the propellant in the combustion chamber. Regenerative cooling in bipropellant engines uses either the fuel or oxidizer as the cooling fluid. Therefore, the thermal energy absorbed by the coolant is not wasted as it augments the initial energy content of the propellant prior to injection, thereby increasing the exhaust velocity and propulsive performance.
Radiation cooling is typically used in monopropellant thrust chambers, some gas generators and for nozzle exhaust sections. Radiation cooling is a simple, lightweight cooling method, which is commonly employed in low-temperature rocket engines, such as hydrazine (monopropellant) spacecraft maneuver and attitude control systems, where the maximum chamber temperature is only about 650xc2x0 C. Refractory metals such as molybdenum, tantalum, tungsten, and niobium have been used in radiation cooled thrust chambers requiring increased operating temperatures up to 1650xc2x0 C. Refractory metals are, however, difficult to fabricate and some suffer from hydrogen embrittlement degradation; others oxidize readily and thus require protective surface coatings to function reliably; all are weight inefficient due to their very high specific gravities. Radiation cooled thrust chambers generally have to protrude beyond the outer skin of the flight vehicle to permit satisfactory radiative heat rejection.
In general, ceramics have superior high-temperature strength and stiffness, and lower density than metallic materials. The principal disadvantages of ceramics as structural materials are their low failure strain, low fracture toughness and catastrophic brittle failure characteristics. Because of these inherent limitations, monolithic ceramics lack the properties of reliability and durability that are necessary for structural design acceptance. However, the emerging technology of fiber-reinforced ceramics, or ceramic matrix composites is one promising solution for overcoming the reliability and durability problems associated with monolithic ceramics. By incorporating high strength, relatively high modulus fibers into brittle ceramic matrices, combined high strength and high toughness composite materials can be obtained. Successfully manufactured ceramic matrix composites exhibit a high degree of non-linear stress-strain behavior with ultimate strengths, failure strains and fracture toughnesses that are substantially greater than that of the unreinforced matrix.
In order to exploit the benefits of fiber reinforcement in brittle ceramic matrices, it is well recognized that a relatively weak fiber/matrix interfacial bond strength is essential for preventing catastrophic failure from propagating matrix cracks. The interface must provide sufficient fiber/matrix bonding for effective load transfer, but must be weak enough to debond and slip in the wake of matrix cracking, leaving the fibers to bridge the cracks and support the far-field applied load. Fiber-reinforced ceramic matrix composites with very high fiber/matrix interfacial bond strengths (usually the result of chemical interaction during manufacture) exhibit brittle failure characteristics similar to that of unreinforced monolithic ceramics by allowing matrix cracks to freely propagate directly through the reinforcing fibers. Conversely, by reducing the interfacial bond strength, the fiber and matrix are able to debond and slip promoting the arrest and/or diversion of propagating matrix cracks at/or around the reinforcing fiber. Since crack inhibition/fracture toughness enhancement is the primary advantage of fiber-reinforced ceramic matrix composites, properly engineered fiber coating systems are thus key to the structural performance of these materials. Control of interfacial bonding characteristics between the fiber and matrix following manufacture and during service is typically provided by the use of applied fiber coatings.
Fiber-reinforced ceramic matrix composites produced by the chemical vapor infiltration (CVI) process are a particularly promising class of engineered high-temperature structural materials, which are now commercially available. The principal advantage of the CVI process approach for fabricating ceramic matrix composites as compared to other manufacturing methods (e.g., reaction bonding, hot-pressing, melt infiltration, or polymer impregnation/pyrolysis) is the ability to infiltrate and densify geometrically complex, multidirectional fibrous preforms to near-net-shape with a ceramic matrix of high purity and controllable stoichiometry without chemically, thermally or mechanically damaging the relatively fragile reinforcing fibers. In addition, because it is a relatively low temperature manufacturing process, high purity refractory matrix materials can be formed (deposited) at a small fraction of their melting temperature (xcx9cTm/4). Despite the many possible high-temperature ceramic matrix composite material systems, however, the number of practical systems is limited by the currently available reinforcing fibers. To date, the majority of high performance ceramic matrix composites produced have focused primarily on carbon and polymer-derived SiC (Nicalon and Hi-Nicalon) fiber reinforcement and CVI-derived SiC matrices.
Carbon fiber-reinforced silicon carbide (C/SiC) and silicon carbide fiber-reinforced silicon carbide (SiC/SiC) composites produced by CVI have been identified by the aerospace and propulsion communities (both in the United States and in Europe) as an important enabling materials technology for thermostructural applications demanding high strength and toughness at temperatures to 1650xc2x0 C. The high purity CVI-SiC matrix is not readily attacked by either hydrogen-rich or oxidizing environments up to 1650xc2x0 C., and resists oxidation by the formation of an adherent and protective oxide surface scale. Along with being chemically compatible with a variety of commercially available refractory fiber reinforcements, SiC possess excellent thermal shock resistance due to its combination of very high thermal conductivity (50-100 W/m/K) and low thermal expansivity (4.5 ppm/xc2x0 C.). These thermophysical attributes are especially attractive for advanced rocket propulsion thrust chamber and exhaust nozzle applications.
Uncooled, single-walled rocket thrust chambers and nozzle components have been produced from fiber-reinforced ceramic matrix composites by a number of manufacturers. Material systems for these applications have included carbon fiber-reinforced carbon, or carbon/carbon (C/C), C/SiC and SiC/SiC composites produced by various manufacturing techniques, including CVI. In all cases, these rocket propulsion devices were fabricated as a simple, single-wall shell construction and passively cooled by radiation.
Until now, actively-cooled reinforced ceramic matrix composite thrust chambers have neither been designed nor produced for various technical reasons, including the lack of conception of a practical design approach, the perceived high level of complexity for such a fiber-reinforced composite design, manufacturing difficulty, and high cost necessary to produce such a device.
In the present invention, a process is described for the design and manufacture of actively-cooled fiber-reinforced ceramic matrix composite tubular shell structures for high-temperature applications with emphasis on converging-diverging thrust chambers for liquid rocket propulsion systems.
Tubular shell structures are defined as any open-ended three-dimensional body with a central longitudinal axis wherein the body of the structure is enclosed by either curved surfaces, flat surfaces or combinations thereof.
Fiber-reinforcement is defined as any refractory fibers, either continuous or discontinuous, used for producing a fibrous preform texture, which are capable of withstanding a use temperature of at least 800xc2x0 C. in an atmosphere which is thermochemically compatible with that fiber without suffering fundamental chemical, physical or mechanical degradation. Examples include carbon fibers, silicon carbide fibers, silicon nitride fibers, aluminum oxide fibers, etc.
A fiber preform is a fibrous texture defined as any assemblage of one or more reinforcing fiber types produced by weaving, braiding, filament winding, fiber placement, felting, needling, or other textile fabrication process.
Fiber preforming is a textile fabrication process by which the collimated multifilamentary fiber bundles (tows) are placed and maintained in a fixed position for purposes of controlling both their orientation and content within a given volumetric space. As such, the spatial arrangement of fibers is referred to as a preform architecture.
The braided architecture is one of the simplest and lowest cost fiber preforms for producing continuous fiber-reinforced axisymmetric, tubular structures. Although constructed of interlaced tows in a planar 2-dimensional arrangement similar to woven fabric, braiding offers the added flexibility of interlacing fiber in three (3) directions both axial and helical. The most common braid architecture consists of helical fibers interlaced at a prescribed bias angle and are termed biaxial braids. Fixed axial fibers can be inserted around the mandrel circumference in nearly any desired fraction with respect to the helical xe2x80x9cbraidersxe2x80x9d to produce an architecture reinforced in three (3) discrete directions and are termed triaxial braids. Triaxially braided preforms offer certain benefits over biaxial architectures, such as the ability to tailor material isotropy or increase axial properties. In most cases, increased axial properties are gained at the cost of sacrificed circumferential xe2x80x9choopxe2x80x9d properties. However, the optimum braid architecture for a given application is usually designed and selected on the basis of the combined axial and hoop properties required. Triaxial preforms typically yield preforms with slightly lower fiber volume fractions than that of the biaxial braided architecture. This is due to the crimping and bunching of the fiber tows at the braid triple point, which results in increased braid layer thicknesses and larger internal voids.
Fiber coating is defined as any refractory composition of either carbon, metal carbide, metal nitride, metal boride, metal silicide, metal oxide, or combinations thereof which is (are) deposited (for example by chemical vapor infiltration) onto the refractory fibers either before or after fiber preforming for purposes of controlling the fiber/matrix interfacial bonding characteristics in the resultant composite. The resultant fiber coating thus encapsulates the reinforcing fibers. Examples include pyrolytic carbon, silicon carbide, silicon nitride, boron carbide, boron nitride, etc.; either as a single-layer phase, multilayered phase or as a phase of mixed composition.
Ceramic matrix is defined as any refractory composition of either carbon, metal carbide, metal nitride, metal boride, metal silicide, metal oxide, or combinations thereof which is subsequently deposited (for example by chemical vapor infiltration) onto the previously coated refractory fibers within the fibrous preform thereby encapsulating the fibers and consolidating the preform into the resultant densified composite. The reinforcing fibers of the fibrous preform thus become embedded within and supported by the surrounding matrix. Examples include pyrolytic carbon, silicon carbide, silicon nitride, boron carbide, boron silicide, etc., either as a single phase, multilayered phase or as a phase of mixed composition.